EP3037570B1 - Herstellungsverfahren einer versiegelungsbeschichtung - Google Patents

Herstellungsverfahren einer versiegelungsbeschichtung Download PDF

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Publication number
EP3037570B1
EP3037570B1 EP15189695.8A EP15189695A EP3037570B1 EP 3037570 B1 EP3037570 B1 EP 3037570B1 EP 15189695 A EP15189695 A EP 15189695A EP 3037570 B1 EP3037570 B1 EP 3037570B1
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EP
European Patent Office
Prior art keywords
topcoat
protrusions
turbine
coating
seal
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Active
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EP15189695.8A
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English (en)
French (fr)
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EP3037570A1 (de
Inventor
John R. Farris
Michael G. Mccaffrey
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RTX Corp
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United Technologies Corp
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    • C—CHEMISTRY; METALLURGY
    • C23—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; CHEMICAL SURFACE TREATMENT; DIFFUSION TREATMENT OF METALLIC MATERIAL; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL; INHIBITING CORROSION OF METALLIC MATERIAL OR INCRUSTATION IN GENERAL
    • C23C—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; SURFACE TREATMENT OF METALLIC MATERIAL BY DIFFUSION INTO THE SURFACE, BY CHEMICAL CONVERSION OR SUBSTITUTION; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL
    • C23C4/00—Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge
    • C23C4/01—Selective coating, e.g. pattern coating, without pre-treatment of the material to be coated
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B05—SPRAYING OR ATOMISING IN GENERAL; APPLYING FLUENT MATERIALS TO SURFACES, IN GENERAL
    • B05D—PROCESSES FOR APPLYING FLUENT MATERIALS TO SURFACES, IN GENERAL
    • B05D3/00—Pretreatment of surfaces to which liquids or other fluent materials are to be applied; After-treatment of applied coatings, e.g. intermediate treating of an applied coating preparatory to subsequent applications of liquids or other fluent materials
    • B05D3/007—After-treatment
    • C—CHEMISTRY; METALLURGY
    • C23—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; CHEMICAL SURFACE TREATMENT; DIFFUSION TREATMENT OF METALLIC MATERIAL; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL; INHIBITING CORROSION OF METALLIC MATERIAL OR INCRUSTATION IN GENERAL
    • C23C—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; SURFACE TREATMENT OF METALLIC MATERIAL BY DIFFUSION INTO THE SURFACE, BY CHEMICAL CONVERSION OR SUBSTITUTION; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL
    • C23C28/00—Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of groups C23C2/00 - C23C26/00 or by combinations of methods provided for in subclasses C23C and C25C or C25D
    • C23C28/04—Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of groups C23C2/00 - C23C26/00 or by combinations of methods provided for in subclasses C23C and C25C or C25D only coatings of inorganic non-metallic material
    • C23C28/042—Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of groups C23C2/00 - C23C26/00 or by combinations of methods provided for in subclasses C23C and C25C or C25D only coatings of inorganic non-metallic material including a refractory ceramic layer, e.g. refractory metal oxides, ZrO2, rare earth oxides
    • C—CHEMISTRY; METALLURGY
    • C23—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; CHEMICAL SURFACE TREATMENT; DIFFUSION TREATMENT OF METALLIC MATERIAL; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL; INHIBITING CORROSION OF METALLIC MATERIAL OR INCRUSTATION IN GENERAL
    • C23C—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; SURFACE TREATMENT OF METALLIC MATERIAL BY DIFFUSION INTO THE SURFACE, BY CHEMICAL CONVERSION OR SUBSTITUTION; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL
    • C23C4/00—Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge
    • C23C4/04—Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge characterised by the coating material
    • C23C4/10—Oxides, borides, carbides, nitrides or silicides; Mixtures thereof
    • C23C4/11—Oxides
    • C—CHEMISTRY; METALLURGY
    • C23—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; CHEMICAL SURFACE TREATMENT; DIFFUSION TREATMENT OF METALLIC MATERIAL; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL; INHIBITING CORROSION OF METALLIC MATERIAL OR INCRUSTATION IN GENERAL
    • C23C—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; SURFACE TREATMENT OF METALLIC MATERIAL BY DIFFUSION INTO THE SURFACE, BY CHEMICAL CONVERSION OR SUBSTITUTION; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL
    • C23C4/00—Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge
    • C23C4/18—After-treatment
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • F01D11/122—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with erodable or abradable material
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/005—Selecting particular materials
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/11—Shroud seal segments
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/29—Three-dimensional machined; miscellaneous
    • F05D2250/294—Three-dimensional machined; miscellaneous grooved
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/20—Oxide or non-oxide ceramics
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present disclosure relates generally to gas turbine engines and, more particularly, to turbine blade outer air seals (BOAS). Specifically, the disclosure concerns coatings applied to turbine blade sealing structures.
  • BOAS turbine blade outer air seals
  • the distance between the tip of the blades and the surface of the BOAS greatly impacts turbine efficiency. Accordingly, efforts have been made to reduce the distance between the blade tips and the BOAS as much as possible.
  • gas turbine engines allow the blade to rub up against a ceramic-coated BOAS during operation, creating wear on either the blade tip, the BOAS surface, or both. With a thermally insulating ceramic coating on the BOAS, the blade usually takes most of the wear.
  • the present invention provides a method of forming a coating as defined in claim 1.
  • FIG. 1 is a simplified cross-sectional view of an exemplary gas turbine engine 10 in accordance with embodiments of the present disclosure.
  • Turbine engine 10 includes fan 12 positioned in bypass duct 14.
  • Turbine engine 10 also includes compressor section 16, combustor (or combustors) 18, and turbine section 20 arranged in a flow series with upstream inlet 22 and downstream exhaust 24.
  • Compressor 16 includes stages of compressor vanes 26 and blades 28 arranged in low pressure compressor (LPC) section 30 and high pressure compressor (HPC) section 32.
  • Turbine section 20 includes stages of turbine vanes 34 and turbine blades 36 arranged in high pressure turbine (HPT) section 38 and low pressure turbine (LPT) section 40.
  • HPT section 38 is coupled to HPC section 32 via HPT shaft 42, forming the high pressure spool.
  • LPT section 40 is coupled to LPC section 30 and fan 12 via LPT shaft 44, forming the low pressure spool.
  • HPT shaft 42 and LPT shaft 44 are typically coaxially mounted, with the high and low pressure spools independently rotating about turbine axis (centerline) C L .
  • Fan 12 includes a number of fan airfoils circumferentially arranged around a rotating member, which is coupled to LPC section 30 and driven by LPT shaft 44.
  • fan 12 is coupled to the fan spool via geared fan drive mechanism 46, providing independent fan speed control.
  • fan 12 is forward-mounted and provides thrust by accelerating flow downstream through bypass duct 14, for example in a high-bypass configuration suitable for commercial and regional jet aircraft operations.
  • fan 12 is an unducted fan or propeller assembly, in either a forward-or aft-mounted configuration.
  • turbine engine 10 includes any of a high-bypass turbofan, a low-bypass turbofan, or a turboprop engine, and the number of spools and shaft configurations may vary.
  • incoming airflow F I enters inlet 22 and divides into core flow F C and bypass flow F B , downstream of fan 12.
  • Core flow F C continues along the core flowpath through compressor section 16, combustor 18, and turbine section 20, and bypass flow F B proceeds along the bypass flowpath through bypass duct 14.
  • LPC section 30 and HPC section 32 of compressor 16 compress incoming air for combustor 18, where fuel is introduced, mixed with air, and ignited to produce hot combustion gas.
  • fan 12 can also provide some degree of compression to core flow F C , and LPC section 30 may be omitted.
  • an additional intermediate spool can be included, for example in a three-spool turboprop or turbofan configuration.
  • thermodynamic efficiency of turbine engine 10 is tied to the overall pressure ratio, as defined between the delivery pressure at inlet 22 and the compressed air pressure entering combustor 18 from compressor section 16. In general, a higher pressure ratio offers increased efficiency and improved performance. It will be appreciated that various other types of turbine engines can be used in accordance with the embodiments of the present disclosure.
  • FIG. 2 is a simplified cross-sectional view along the line 2-2 of FIG.1 of casing 48 having rotor shaft 50. Blades 36 are attached to rotor shaft 50, and the gas path is shown as the space between blades 36. As shown in FIG. 3 , topcoat 52, corresponding to the coating of this disclosure, is on substrate 56, such that the clearance C between topcoat 52 (topcoat 52 includes insulating layers 58)and blade tips 54 of blades 36 serves as a seal to prevent leakage of air, thus improving engine efficiency. In FIG. 2 , clearance C is enlarged for the purposes of illustration.
  • clearance C can be, for example, in a range of about 0.025 inches (0.064 centimeters) to 0.055 inches (0.14 centimeters) when the engine is cold and 0.000 to 0.035 inches (0.09 centimeters) during engine operation, depending on the specific operating conditions and previous rub events that may have occurred.
  • FIG. 3 is a cross-sectional view of HPT section 38 along line 3-3 of FIG. 2 with blade tip 54 of blade 36 and topcoat 52.
  • Topcoat 52 is attached to casing 48 via substrate 56, resulting in a clearance C between topcoat 52 and blade tip 54 of blade 36 that varies with operating conditions, as described herein.
  • Substrate 56 can be any material suitable for forming a desired geometry between casing 48 and thermally insulating layers 58.
  • substrate 56 can be formed from a metal alloy, such as a nickel-based alloy, and affixed to casing 48 by welding or other suitable method.
  • substrate 56 can be additively manufactured to form the desired pattern.
  • a bond coat (not shown) can be used to facilitate bonding between substrate 56 and thermally insulating layers 58.
  • the bond coat can include any type of bonding material suitable for attaching substrate 56 and thermally insulating layers 58.
  • Substrate 56 includes protrusions 60, which can be geometric surface features on substrate 56.
  • Protrusions 60 extend into the gas flow path such that gaps are formed between adjacent protrusions.
  • Protrusions 60 can be of substantially uniform height.
  • the height of protrusions 60 can be at least 0.01 inches (0.025 centimeters), such that the gaps between protrusions 60 have some depth greater than the desired clearance C.
  • Topcoat 52 includes one or more thermally insulating layers 58 layered over protrusions 60.
  • topcoat 52 includes two thermally insulating layers 58.
  • thermally insulated layers 58 can include any number of layers formed from any number of materials.
  • Thermally insulating layers 58 can be formed from any abradable material suitable for providing a desired heat resistance within the gas path such as ceramic or metallic materials.
  • the material can be a ceramic composite such as yytria-stabilized zirconia, hafnia, and/or gadolinia, gadolinia zirconate, molybdate, alumina, or combinations thereof.
  • Thermally insulating layers 58 can include various porosities.
  • Thermally insulating layers 58 can be deposited on substrate 56 by any number of thermal spray processes, such as plasma spray, electron beam, high velocity oxygen fuel (HVOF), or cathodic arc. Thermally insulating layers 58 are in an unconverted, intermediate state when applied to substrate 56.
  • thermal spray processes such as plasma spray, electron beam, high velocity oxygen fuel (HVOF), or cathodic arc. Thermally insulating layers 58 are in an unconverted, intermediate state when applied to substrate 56.
  • Protrusions 60 produce faults 62 in topcoat 52.
  • the faults can be produced using any variety of different protrusions 60.
  • protrusions 60 can be cylindrical, rectangular, triangular, or any other three-dimensional shape.
  • Protrusions 60 can have substantially uniform height, preferably at least 0.01 inches (0.25 centimeters).
  • protrusions 60 form a grid pattern.
  • protrusions 60 form other patterns to create the desired faults 62 in topcoat 52.
  • topcoat 52 can have any surface pattern with dimples 64 formed between protrusions 60. Faults 62, running through topcoat 52 from protrusions 60, are bounded by segmented portions in topcoat 52.
  • segmented portions serve as separation points in thermally insulating layers 58 that reduce the transmission of thermally induced stresses from one region to another in topcoat 52, and further reduce the effect of the thermal and mechanical stresses from blade tip 54 during an initial break-in cycle (discussed in FIG. 4A ).
  • FIG. 4A is a cross-sectional view of HPT section 38 of FIG. 3 during a first phase of a process to prepare an abradable, durable coating on substrate 56.
  • FIG. 4B is a cross-sectional view of HPT section 38 after the first phase of the process.
  • Blade 36 of HPT section 38 is enclosed in casing 48.
  • Substrate 56 is attached to casing 48.
  • Substrate 56 includes protrusions 60, which extend outward from substrate 56 toward blade tip 54 of blade 36.
  • Topcoat 52 is deposited on substrate 56.
  • Topcoat 52 includes thermally insulating layers 58 layered over protrusions 60.
  • Protrusions 60 form faults 62 extending from substrate 56 through topcoat 52.
  • turbine engine 10 (not shown) is undergoing the first phase of a break-in cycle or green run.
  • abradable coatings of a blade outer air seal are made from compositions that are easily worn down by the passing blade tip. These abradable coatings remain soft enough to allow the blade tip to cut into the coating during normal engine operation. In this manner, some of the wear that typically occurs to the blades is shared by the BOAS. Over time, however, the softer, abradable, coatings erode due to particles entering the turbine gas path. This erosion creates a larger clearance between the blade tip and BOAS. The larger the clearance between the blade tip and the BOAS, the less efficient the engine becomes.
  • BOAS blade outer air seal
  • the BOAS coating hardness is increased to resist particle erosion, it typically becomes too hard for the blade tip to effectively wear the coating. As a result, the blade tip strikes the hard coating and melts due to friction. Molten material from the blade tips can then collect and build up on the BOAS coating surface, forming a dam or wall that further wears the blade tip.
  • topcoat 52 is installed in the turbine in a soft, unconverted state.
  • the engine is rapidly accelerated. That is, the engine is run up to its maximum power in a very short time.
  • This rapid acceleration causes blade tip 54 of blade 36 to rub topcoat 52.
  • This initial rub creates channel 66 (described below in greater detail and shown in FIG. 4B ) in topcoat 52 that corresponds to the wear pattern blade tip 54 will create in topcoat 52 over time during normal operation.
  • topcoat 52 is cut to depth D (described below in greater detail and shown in FIG. 4B ) by blade tip 54 while remaining attached to substrate 56.
  • Protrusions 60 create greater surface area between substrate 56 and topcoat 52 to hold topcoat 52 in place during rub events. Depth D is less than the depth of the dimples 64 such that dimples 64 are not worn away by the initial rub.
  • protrusions 60 can cause topcoat 52 to be more easily cut in a desired pattern during the initial rub phase.
  • protrusions 60 are substantially the same height, and substantially evenly spaced. In this arrangement, protrusions 60 form a dimpled surface when thermally insulating layers 58 are deposited over substrate 56.
  • protrusions 60 can take on any pattern that creates a desired fault pattern in topcoat 52. In this manner, a desired wear pattern can be formed in topcoat 52 after the BOAS has been installed in the turbine during a break-in cycle or green run.
  • turbine engine 10 is undergoing the second phase of the break-in cycle. If topcoat 52 remained in its unconverted state after the first phase, it would be continually worn down by erosion. Thus, during the second phase of the break-in cycle, the surface of topcoat 52 is converted by a high-temperature event that hardens topcoat 52, converting topcoat 52 into a durable surface while preserving channel 66. Channel 66 can thereby retain a channel depth D for longer over the life of the engine because the converted coating is more resistant to particle erosion.
  • topcoat 52 is run through a two-phase break-in process to create a desired wear pattern in a hardened surface for optimal sealing between blade tip 54 and topcoat 52.
  • FIG. 5 is a flow diagram illustrating a process (110) of preparing an abradable, durable coating with continuing reference to topcoat 52 of FIGS. 3 , 4A , and 4B .
  • Preparing an abradable, durable topcoat 52 for a BOAS includes forming substrate 56 (step 112), depositing thermally insulating layers 58 to form coating (step 114), installing coated BOAS in casing (step 116), rubbing blade tip 54 against topcoat 52 (step 118), and converting topcoat 52 (step 120).
  • Step 112 includes forming substrate 56 as described in FIG. 3 , with protrusions 60 extending from substrate 56.
  • Step 114 follows step 112 and includes depositing thermally insulating layers 58 to form topcoat 52.
  • Substrate 56 can be a metal alloy or additively manufactured material as discussed in FIG. 3 .
  • Thermally insulating layers 58 are deposited over protrusions 60 by thermal spray or other deposition methods to form topcoat 52, which is in an unconverted, intermediate state.
  • the coated BOAS is installed in casing 48 of turbine engine 10.
  • Step 118 follows step 116, and includes rubbing blade tip 54 against topcoat 52.
  • topcoat 52 has a high degree of abradability. Rapid acceleration during step 118 creates an initial rub that can produce equal wear to blade tip 54 and topcoat 52.
  • turbine engine 10 is run at a steady state high temperature to cause conversion of topcoat 52. Topcoat 52 stays attached to substrate 56 when converted due to protrusions 60 extending from substrate 56. Thus, channel depth D is fixed in topcoat 52. In this manner, the equal wear to blade tip 54 and topcoat 52 is preserved over the life of the engine.
  • an abradable, durable coating can be formed to provide long-lasting, effective sealing between blade tip 54 and casing 48.
  • the seal may be a blade outer air seal.
  • the topcoat may comprise a thermally insulating ceramic material having a porosity between 5 and 70 volume percent.
  • converting the topcoat may comprise running an engine to expose the seal to a surface temperature of at least 1066 °C (1950 °F).

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  • Chemical & Material Sciences (AREA)
  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Materials Engineering (AREA)
  • Chemical Kinetics & Catalysis (AREA)
  • Metallurgy (AREA)
  • Organic Chemistry (AREA)
  • Plasma & Fusion (AREA)
  • Physics & Mathematics (AREA)
  • General Engineering & Computer Science (AREA)
  • Ceramic Engineering (AREA)
  • Inorganic Chemistry (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Claims (5)

  1. Verfahren zum Herstellen einer Beschichtung, umfassend:
    Abscheiden eines Substrats (56), das eine Vielzahl von Vorsprüngen (60) aufweist, auf eine Dichtung;
    Abscheiden einer Deckschicht (52) in einem weichen, nicht umgewandelten Zustand über die Vorsprünge (60), wobei die Deckschicht (52) segmentierte Abschnitte umfasst, die eine Vielzahl von Brüchen (62) definieren, die sich von der Vielzahl von Vorsprüngen (60) durch die Deckschicht (52) erstrecken, wobei das Abscheiden der Deckschicht (52) über die Vielzahl von Vorsprüngen (60) eine Fläche bildet, die eine Vielzahl von Vertiefungen (64) aufweist,
    gekennzeichnet durch:
    Erzeugen eines Verschleißmusters eines Kanals (66) in der Deckschicht (52) während eines anfänglichen Reibereignisses durch schnelles Beschleunigen der Turbine während eines ersten Laufs durch Laufenlassen der Turbine bis zu einer maximalen Leistung in einer sehr kurzen Zeit, was das Reiben der Schaufelspitze (54) gegen die Beschichtung (52) bewirkt, um den Kanal (66) zu bilden, der eine Tiefe aufweist, die geringer als eine Tiefe der Vielzahl von Vertiefungen (64) ist; und dann
    Umwandeln der Deckschicht (52) während eines Hochtemperaturereignisses, wobei das Umwandeln der Deckschicht (52) das Erhalten des Kanals (66) umfasst, der während des anfänglichen Reibereignisses erzeugt wurde.
  2. Verfahren nach Anspruch 1, wobei die Dichtung eine äußere Laufschaufel-Luftdichtung ist.
  3. Verfahren nach Anspruch 1 oder 2, wobei die Deckschicht (52) ein thermisch isolierendes keramisches Material umfasst, das eine Porosität zwischen 5 und 70 Volumenprozent aufweist.
  4. Verfahren nach einem der vorstehenden Ansprüche, wobei die Deckschicht (52) mindestens eine thermisch isolierende Schicht (58) umfasst, die aus einem abreibbaren Material eines keramischen oder metallischen Materials gebildet ist, wobei das keramische Material ein Keramikverbundwerkstoff ist, der Yttrium-stabilisiertes Zirconiumdioxid, Hafniumoxid und/oder Gadoliniumoxid, Gadoliniumoxid-Zirconat, Molybdat, Aluminiumdioxid oder Kombinationen davon umfasst.
  5. Verfahren nach einem der vorstehenden Ansprüche, wobei das Umwandeln der Deckschicht (52) das Laufenlassen eines Triebwerks, um die Dichtung einer Oberflächentemperatur von mindestens 1066 °C (1950 °F) auszusetzen, umfasst.
EP15189695.8A 2014-12-15 2015-10-14 Herstellungsverfahren einer versiegelungsbeschichtung Active EP3037570B1 (de)

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US10823412B2 (en) 2017-04-03 2020-11-03 Raytheon Technologies Corporation Panel surface pockets for coating retention
US11352890B2 (en) * 2017-06-12 2022-06-07 Raytheon Technologies Corporation Hybrid thermal barrier coating
US11015474B2 (en) 2018-10-19 2021-05-25 Raytheon Technologies Corporation Geometrically segmented abradable ceramic thermal barrier coating with improved spallation resistance
US10927695B2 (en) 2018-11-27 2021-02-23 Raytheon Technologies Corporation Abradable coating for grooved BOAS
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US11702950B2 (en) 2023-07-18
US20160201498A1 (en) 2016-07-14
US20200025014A1 (en) 2020-01-23

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